Gas turbine engine

ABSTRACT

A gas turbine engine has a variable pitch fan and a first compressor mounted on one shaft and driven from a common turbine, together with a high pressure compressor and turbine on a second shaft, ducting to supply bleed air to high lift devices on an aircraft in which the engine is mounted, and means for bleeding air from the first compressor to the high lift devices when the fan blades are turned to fine pitch. A control system is included to maintain the first compressor speed at maximum when the high lift devices are being used.

D United States Patent 1191 1111 3,761,042 Denning I Sept. 25, 1973 YGAS TURBINE ENGINE 3,255,585 6/1966 Grieb 244 12 D 3,472,321 10 19691311' 60 226 R [75] Inventor Ralph Mmch Denning Bnstol 3 503 572 3/1970Ha fiiii 244/42 R x England 3,638,421 2/1972 Chilman.... 60/226 R x2,873,576 2/1959 Lombard 60/239 [73] Asslgnee' The f f E Defence3,123,322 3/1964 Jackson 6131.... 244/42 cc M s Gowmmem 3,255,585 6/1966Grieb 244/12 D of P 3,356,152 12/1967 Accorsi 416/27 Brim" and NorthernIreland, 3,604,662 9/1971 Nelson et a1. 244/53 R London, England 22 F, dM 4 1971 FOREIGN PATENTS OR APPLICATIONS 1 ay 502,548 5/1954 Canada416/27 [21] App]. No.: 140,235 881,662 11/1961 Great Britain 244/12 [30]Foreign Application Priority Data m Exami' wr Miltn Buchler Ma l 6 1970G t B 23 839/70 Assistant ExammerBarry L. Keimachter y rea n n July 1,1970 Great Britain 31,937/70 Attorney Davls M1 [er & Mosh [52] US. Cl244/53 R, 60/226 R, 244/12 D, [57] ABSTRACT 244/23 D A gas turbineengine has a variable pitch fan and a first compressor mounted on oneshaft and driven o a Fleld of Search R, CC, common turbine, together a psu e c0 244/23 42 60/226 416/27 pressor and turbine on a second shaft,ducting to supply I bleed air to high lift devices on an aircraft inwhich the [5 6] References cued engine is mounted, and means forbleeding air from the UNITED STATES PATENTS first compressor to the highlift devices when the fan 2,873,576 2/1959 Lombard 60/239 blades areturned to fine pitch. A control system is in- 3,008,S15 1 1/196! Wente60/243 cluded to maintain the first compressor speed at maxi- 3,054,2549/1962 pp 60/243 X mum when the high lift devices are being used.3,108,767 10/1963 E1tis et a1 60/226 X 3,123,322 3/1964 Jackson et aL.244/42 CC 7 Claims, 5 Drawing Figures 3,191,886 6/1965 Lewis et 3160/226 R,

PATENTED SEP25IH73 SHEET 1 0F 4 FIG I PAIENTEDSEPZSIQTS ISO SHEET 3 BF 4I A MAX. NH LIMIT TI-IRUST INCREASING FIG. 4

FIG. 3

GAS TURBINE ENGINE The present invention relates to gas turbine enginesand relates in particular to gas turbine engines which are adapted toprovide a bleed of air for use in operating high lift devices on anaircraft.

It has been proposed in our Patent Specification No. 1,127,853 to adapta three-shaft engine for providing the required air bleed. In thatspecification we provide a fixed pitch front fan driven by a lowpressure (LP) turbine and mounted on its own L.P. shaft. Part of the aircompressed by the fan provides forward thrust through a forwardpropulsion nozzle while theremainder of the air passes into anintermediate pressure (IP) compressor driven by an LP. turbine via anLP. shaft. The flow through the LP. compressor is split from a pointabout half-way along its length, part of the flow passing into theremaining stages of the LP. compressor and hence to the high pressure(HP) compressor, while the remainder ofthe flow passes into a duct whichdirects it either to the high lift device or to an additional forwardpropulsion nozzle.

One disadvantage of an engine as described above is that when theaircraft is making its approach run the forward thrust required isminimum while the bleed to the high lift device is required to be amaximum, and these two conditions are incompatible in the abovearrangement.

It is the object of the present invention to provide a gas turbineengine which overcomes this disadvantage.

According to the present invention, there is provided a gas turbineengine for use in an aircraft having a fluid operated high lift device,the engine comprising a first compressor and a second compressor in flowseries, a fan drivingly connected to the first compressor and means forbleeding working fluid from the first compressor to the high lift devicecharacterized in that the second compressor is driven from an H.P.turbine by means of a first shaft, the fan and the first compressor aredriven from a common lower pressure turbine by means of a second shaft,the fan is a variable pitch fan, and control means are provided forcontrolling the speeds of the two shafts when the pitch of the fan isreduced.

The fan may be a front fan or an aft fan.

In one form of engine according to the invention the bleed to thefluid-operated high lift devices is used only when the aircraft is onits landing approach to an airfield. On take-off the fan blading is setto a pitch required for maximum fan thrust and all the air compressed bythe first compressor passes into the second compressor.

In another form of engine, however, the first compressor is dimensionedto compress an amount of air which is surplus to that required for thesecond compressor and a permanent bleed is taken from downstream of theintermediate pressure compressor.

According to a feature of the present invention therefore, there isprovided a gas turbine engine for use in an aircraft having one or moreair operated high lift devices, the engine comprising a front fan, anLP. compressor, and an HP. compressor all in flow series, the LP.compressor being adapted to compress an amount of air surplus to thatrequired by the H.P. compressor, and ducting along which said surplusair may pass to the high lift device or devices, characterized in thatthe fan and the LP. compressor are driven from a common LP. turbine bymeans of a first shaft, the l-I.P. compressor is driven from an HP.turbine by means of a second shaft, the fan is a variable pitch fan, anda governor is provided which limits the maximum speed of the first shaftwhen the pitch of the fan is reduced, by controlling the fuel suppliedto the engine.

In a preferred embodiment of the invention the ducting along which saidsurplus air passes is formed with a secondary passage which terminatesin a rearwardly pointing propulsion nozzle, and valve means are providedfor selectively directing the air to the high lift device or devices,and the propulsion nozzle.

With engines constructed according to the present invention, in theparticular case of the aircraft ap proaching an airfield, the forwardthrust of the engine can be reduced, while the air supply to the highlift device or devices is maintained at a high level by turning the fanblades to a fine pitch setting.

The term variable pitch fan as used in this specification is to be takento include a fan. the blades of which are rotatable about theirlongitudinal axes to vary their angle of attack.

The high lift devices in the preferred embodiments of the invention arethe flaps in the aircraft wings over or through which said surplus airpasses to produce additional lift on the wings.

The fan may be directly connected to the shaft of the first compressoror may be connected thereto through a gearbox.

One example of the invention will now be more par ticularly describedwith reference to the accompanying drawings in which FIG. 1 illustratesan aircraft incorporating the present invention and,

FIG. 2 is an enlarged view which illustrates diagrammatically the layoutof one of the engines of the aircraft of FIG. 1.

FIG. 3 illustrates diagrammatically a flow diagram for the control meanswhich controls the speeds of the shaft.

FIG. 4 is a graph showing generally the relationships between shaftspeeds for various .pitch settings and thrust levels, and

FIG. 5 illustrates an alternative engine layout which includes areduction gear box.

Referring now to the drawings FIG. 1 shows an aircraft 1 underneath thewings 3 of which, four engines 2 are mounted on pylons 4. Along each ofthe wings are flaps 5 for aircraft control and a duct 6 extends alongthe wing adjacent the. flaps to carry compressed air which is ejectedover the flaps from slit nozzles (not shown) along the rearward side ofthe duct. The air is supplied to the duct 6 from each engine via ducting7 in the pylon.

Turning now to FIG. 2 one of the engines 2 is shown in a cut-away viewwhich shows the top half only of the engine.

The engine comprises a front fan 10 mounted for rotation in a duct 1 1which terminates downstream of the fan in a propulsion nozzle 12. Partof the air compressed by the fan passes rearwards, radially inwardly ofthe duct, into an LP. compressor 14. The LP. compressor is made largeenough to compress an amount of air surplus to the requirements of anHP. compressor l6 downstream thereof, so that only part of the aircompressed by the LP. compressor passes rearwardly into the HP.compressor and hence to the combustion equipment 18.

The hot gases produced by the combustion equipment pass rearwardly intoan HP. turbine 20 which drives the l-l.P. compressor by means of an HP.shaft 22, and then to an LP. turbine 23 which drives both the LP.compressor 14 and the fan by means of an L.P. shaft 24. The hot gasefflux from the L.P. turbine finally passes to atmosphere through a hotgas propulsion nozzle 25.

The H.P. and L.P. shafts are supported for rotation in bearingsrepresented at 26, 27, 28 and 29 respectively.

The surplus air compressed by the [.P. compressor passes into theducting 7, in the pylon 4, which branches into two passages, one leadingto the duct 6 in the wing for supplying the air to the aircraft flaps 5,and the other leading to a propulsion nozzle 32. Valves 33 and 34 aredisposed in the ducting 7 for selectively closing one of the passageswhile opening the other.

Control of the valves is carried out directly by a pilots lever in theaircraft cockpit.

The fan 10 is provided with a pitch varying mechanism 36 which can be ofany suitable form, known per se, and is not described in detail. Themechanism 36 is under the control of a pilot's lever either directly orvia the engine fuel system 41, e.g., the pilot may select H.P. shaftrpm. and the fuel system automatically controls the fan pitch from this.

The ratio of the thrust split between the fan propulsion nozzle, thebleed air, and the hot gas propulsion nozzle, may be of the order of 40percent fan thrust, 40 percent bleed thrust'and 20 percent hot gasthrust.

In operation such an engine would have a low hot jet velocity and wouldberelatively quiet. On take-off the pilot selects the coarse pitchsetting for the fan; valve 33 to be open and valve 34 closed, thusproviding 40 percent bleed thrust for the high lift devices to produce ashort take-off. The air exhausting from the flaps gives some propulsiveforward thrust, which is added to the thrust from the fan and hot gasnozzles.

Once in the air, the pilot changes over the selection of valves 33 and34 so that valve 33 is closed and valve 34 open. This gives the normalcrusing configuration and thrust is provided from nozzle 32.

On approach to an airfield the pilot again selects valve 33 to be openand valve 34 closed but now in addition the pilot also selects a finerpitch for the fan.

The effect of this is to take a significant load off the LP. turbine andhence it tends to overspeed. The LP. compressor which is also driven bythe LP. turbine also overspeeds, and an engine control governor (seeFIG. 3), acts on the engine fuel system to reduce the engine speed tobring the L.P. turbine back to its governed speed. This in turn causesthe engine governor to reduce the speed of the HP. shaft and thisconsequently reduces hot gas thrust from the propulsion nozzle 25.

The effect is therefore, that the fan thrust and the hot gas thrust areboth reduced by an amount depending on the degree by which the fan pitchis reduced, while the LP. compressor continues to run at its maximumspeed to provide air for flap blowing. There will, of course, be someloss of pressure from the LP. compressor outlet, due to loss of thrustacross the fan which supercharges the LP. compressor. This will becompensated at least in part by the duct 7 receiving a greater portionof air from the LP. compressor due to the l-I.P. compressor running moreslowly.

A simple variable pressure loss producing mechanism, e.g., a butterflyvalve, may be put into the duct to promote matching of the excess l.P.compressor flow to the flow required by the high lift device under theseconditions.

FIG. 3 shows one example of a control system for the engine describedabove, and FIG. 4 illustrates the relationship between the shaft speedsat different thrust levels and pitch settings.

Referring initially to FIG. 4 Fan speed Nf is plotted vertically, andhigh pressure compressor shaft speed Nl-I is plotted horizontally. Thelimits of maximum fan speed and fine pitch are shown, and the verticaldotted lines indicate different total engine thrust levels, thrustincreasing to the right, i.e., as NH increases. Line A A indicates apitch setting required to give designed thrust at the designed maximumfor speed, i.e., at the design point X with no lift augmentation.

Considering a fixed pitch fan engine, in order to reduce the thrustlevel required for landing approach, say YY, the engine operatingparameters are constrained to move down the constant pitch line AA sothat fan speed, and HP. speed also decrease, and hence the bleed airavailable for the high lift devices also decreases. In an engine of thepresent invention however, in which the pitch is variable, the engineparameters can move along the constant maximum Nfline until the thrustcondition Y Y is reached, or until the fine pitch limit is reached afterwhich the engine speed will have to decrease along the fine pitchlimiting line.

The engine can be designed so that the fine pitch limit and the approachthrust value cross on the maximum fan speed line at point Z.

Hence it can be seen that with the present invention, an increase in fanspeed, and hence I.P. compressor speed, of A Nf, is achieved over afixed pitch fan engine, and bleed air flow is maintained at the reducedthrust level.

Referring now to FIG. 3, one control system which achieves the desiredconditions for various fan pitch settings, includes a connection from apilots lever to a pitch change mechanism operating control 151. Thecontrol 151 actuates the pitch change mechanism of the engine. A fanspeed governor 152 detects the change in fan speed and sends a signal toa fuel supply control unit 153 which in turn acts on the fuel system ofthe engine. The pitch change mechanism operating control 151 includesthe fine pitch limiting stop, and

the governor controls the fan maximum speed.

The fan is designed to run at maximum speed during normal cruising andtake-off conditions, so that when the pitch of the fan is changed tofine pitch the fan tends to overspeed, and the governor maintains themaximum speed by reducing the fuel flow to the engine.

Similar effects can be obtained with a system operated by the pilot'slever sending a signal directly to an HP. shaft speed governor, whichacts on the fuel system to reduce engine speed. The speed reduction ofthe fan shaft can then be detected and a signal sent to the pitch changemechanism to vary the pitch to maintain the fan speed constant.

The system may be designed as an on-ofi system in which two pitchpositions only of the fan are used for flap blowing or cruiseconditions, or the system may be arranged to allow an infinite variationof fan pitch between the cruise position and the fine pitch limitingstop.

In an alternative form of engine the air bleed may only be used when theaircraft is on its approach run for landing on an airfield. In this casethere is no need for the propulsion nozzle 32 and there will be no flapblowing on take-off.

This leads to a simpler engine design. The bleed may be up to percent ofthe normal l.P. compressor mass flow, and the fan pitch is varied asrequired to keep the LP. compressor operating near its designcharacteristic (i.e., to avoid compressor surge and fan blade flutter).

In this way the engine thrust can be reduced and added lift can beprovided on the aircraft during approach, as described in relation tothe engine shown in the figures, but for take-off the fan is turned to acoarse pitch to give maximum fan thrust while the whole of the aircompressed by the LP. compressor passes into the I-I.P. compressor togiven maximum hot gas thrust. Some bleed air may be used on take-off toincrease the aircraft lift, but it is believed that at take-offconditions the most efficient way of using the air available would be tohave no bleed.

Referring now to FIG. 5 there is shown a further alternative enginelayout in which a fan 100 and an LP. compressor 101 are mounted on an1.1. shaft 102 which is driven from a turbine 103. A gas generatorcomprising the usual H.P. compressor 105, combustion 105, combustionequipment 106 and I'I.P. turbine 107 provide the hot gases for drivingthe turbine 103.

The fan is connected to the Ll. compressor through a reduction gearbox 110, and is a variable pitch fan, the pitch changing mechanism beingillustrated at 112.

Part of the air compressed by the fan passes out of a fan nozzle 115,while the remainder passes down an annular inlet passage 116 to the LP.compressor 101.

The LP. compressor is designed to compress an amount of air considerablygreater than that required by the gas generator compressor 105, and theexcess air is continuously fed to a bleed duct 120 which communicateswith several ducts 121, 122 and 123 through a valve chamber 125.

The operation of this engine is very similar to the op eration of theengine described with reference to FIG. 2 in that for take-off andapproach valves in the chamber 125 are opened such as to allow the airfrom duct 120 to pass to the duct 121, from which it passes to the wing3 of the aircraft. The air then passes through slit nozzle 128 in thetrailing edge of the wing to pass between upper and lower flaps 129 and130, to augment the lift of the wing.

During normal cruising flight the air is allowed to pass, by openingdifferent valves in chamber 125 to a propulsion nozzle 134 at the end ofduct 122, and after landing, a further valve may be used to allow theair to pass to duct 123 from which it issues forwardly to provide athrust reversing effect.

The control for this engine may be similar to that described withreference to FIGS. 2, 3 and 4.

The engines of the present invention provide a range of power plantswhich offer a wide choice of operating conditions. Although thepreferred split of thrusts between fan exhaust, bleed air, and hot gasexhaust is of the order of 40 percent fan exhaust, 40 percent bleed airand 20 percent hot gas exhaust a wide range of variations is possible.

For example in an aircraft where lift augmentation is required only onapproach, the engine is designed without an oversize l.P. compressor andrelies on up to 20 percent of the LP. compressor air bled to the highlift device when the fan is turned to fine pitch, as described above.

Where large bleeds are required for aircraft wing lift augmentingdevices, the oversize I.P. compressor can be used as described hereinand the thrust split may be such that up to 60 percent of the enginethrust is produced by the bleed air.

The above examples have been described with particular reference to afront fan engine but it will be understood that the fan may be an aftfan, and the principle of operation would be the same.

In either case this fan may be connected directly to the shaft of theLP. turbines, or may be connected through a reduction gearbox to reducefan r.p.m. and hence fan noise.

I claim:

1. A gas turbine engine for use in an aircraft having a fluid operatedhigh lift device, said engine comprising:

a. a first compressor;

b. a second compressor in flow series with said first compressor; 7

c. a fan drivingly connected to said first compressor,

said fan being a variable pitch fan;

d. means for bleeding fluid from said first compressor to the high liftdevice;

e. alow pressure turbine;

f. a first shaft drivingly connecting said fan and said first compressorto said low pressure turbine;

g. a higher pressure turbine;

h. a second shaft drivingly connecting said second compressor to saidhigher pressure turbine; and

i. control means for controlling the speeds of said first and secondshafts when the pitch of said fan is varied.

.2. A gas turbine engine according to claim 1 wherein said fan is afront fan.

3. A gas turbine engine according to claim 1 further including areduction gear connected between said first compressor and said fan.

4. A gas turbine engine according to claim 1 further including a firstauxiliarly nozzle; and wherein said first compressor is dimensioned tocompress an amount of air which is in surplus of that required for saidsecond compressor; and wherein said means for bleeding includes a valvemeans which is selectively operable for directing air to the high liftdevice or said first auxiliarly propulsion nozzle, and ducting fordirecting the surplus air to said valve means.

5. A gas turbine engine according to claim 1 further including a secondauxiliary propulsion nozzle connected to said valve means, said firstpropulsion nozzle being rearwardly pointing for producing forward thrustand said second propulsion nozzle being forwardly pointing for producingreverse thrust.

6. A gas turbine engine according to claim 1 wherein said control meanscomprises means for detecting the variations in speed of said firstshaft, a fuel supply control means, and means connecting said detectingmeans and said fuel supply control means for varying the fuel supply tothe engine as the fan speed varies.

7. A gas turbine engine according to claim 6 wherein said control meansadditionally comprises a fine pitch limiting stop and said detectingmeans includes a governor which limits the maximum speed of said firstshaft.

* i t t i

1. A gas turbine engine for use in an aircraft having a fluid operatedhigh lift device, said engine comprising: a. a first compressor; b. asecond compressor in flow series with said first compressor; c. a fandrivingly connected to said first compressor, said fan being a variablepitch fan; d. means for bleeding fluid froM said first compressor to thehigh lift device; e. a low pressure turbine; f. a first shaft drivinglyconnecting said fan and said first compressor to said low pressureturbine; g. a higher pressure turbine; h. a second shaft drivinglyconnecting said second compressor to said higher pressure turbine; andi. control means for controlling the speeds of said first and secondshafts when the pitch of said fan is varied.
 2. A gas turbine engineaccording to claim 1 wherein said fan is a front fan.
 3. A gas turbineengine according to claim 1 further including a reduction gear connectedbetween said first compressor and said fan.
 4. A gas turbine engineaccording to claim 1 further including a first auxiliarly nozzle; andwherein said first compressor is dimensioned to compress an amount ofair which is in surplus of that required for said second compressor; andwherein said means for bleeding includes a valve means which isselectively operable for directing air to the high lift device or saidfirst auxiliarly propulsion nozzle, and ducting for directing thesurplus air to said valve means.
 5. A gas turbine engine according toclaim 1 further including a second auxiliary propulsion nozzle connectedto said valve means, said first propulsion nozzle being rearwardlypointing for producing forward thrust and said second propulsion nozzlebeing forwardly pointing for producing reverse thrust.
 6. A gas turbineengine according to claim 1 wherein said control means comprises meansfor detecting the variations in speed of said first shaft, a fuel supplycontrol means, and means connecting said detecting means and said fuelsupply control means for varying the fuel supply to the engine as thefan speed varies.
 7. A gas turbine engine according to claim 6 whereinsaid control means additionally comprises a fine pitch limiting stop ;and said detecting means includes a governor which limits the maximumspeed of said first shaft.